With reference to FIG. 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
The high pressure turbine components of an aero gas turbine engine are located in the hottest part of the engine. At around 1600° C., the temperature of the gas stream is greater than the melting temperature of the nickel-based alloys from which the high pressure turbine nozzle guide vanes and rotor blades are typically cast.
It is usual, therefore, to cool nozzle guide vanes and rotor blades internally using cooling air taken from the exit of the high pressure compressor. The cooling air, which bypasses the combustor, may nonetheless be at a temperature of over 700° C. on exit from the compressor section. However, the cooling air, even if returned into the flow path of the turbine downstream of the combustor, does not return a full measure of work to the operation of the turbine. Thus, the greater the amount of cooling air that is extracted, the greater the associated loss in engine efficiency.
One method of reducing the cooling air requirement is to cool the cooling air before it enters the high pressure turbine components. This can be achieved by putting the cooling air in heat exchange relationship with a cooler fluid. For example, many aero gas turbine engines have a bypass air stream which can serve as the cooler fluid. In conventional arrangements, a portion of that air stream is diverted at an offtake into a duct in which the heat exchanger is located. The diverted portion of air, after having passed over the heat exchanger, is then returned to the main bypass air stream.
U.S. Pat. No. 4,254,618 proposes bleeding cooling air from the compressor discharge of a turbofan engine, and routing the cooling air to a heat exchanger located in a diffuser section of the fan duct. The cooled cooling air is then routed through compressor rear frame struts to an expander nozzle and thence to the turbine.